A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine general includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
It is desirable to operate the gas turbine engine at relatively high temperatures and pressures to increase an efficiency of the gas turbine engine or to maximize a power output of the gas turbine engine. For example, it is especially desirable to operate the gas turbine engine at relatively high temperatures and pressures during modes of operation requiring high power output, such as during takeoff and climb of an aircraft having such an engine. However, in order to withstand such relatively high temperatures and pressures, various components of the gas turbine engine must be constructed of rare and/or expensive materials.
Accordingly, a device for cooling various components of the gas turbine engine without substantially reducing an operating temperature and/or pressure of the gas turbine engine would be useful. More particularly, a device for cooling various components of the gas turbine engine not directly exposed to the core air flowpath of the gas turbine engine would be particularly beneficial.